This disclosure relates to a vaporization cooling arrangement for a gas turbine engine rotor.
A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustor section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines.
A circumferential array of blades is mounted on each rotor provided at various stages in the compressor and turbine sections. The thermal fight between the cold inner diameter of the rotor and the hotter portion near the core flow path can be significant, for example, over a 500° F. temperature differential between the inner and outer rotor portions. This temperature differential leads to a complex array of structural failure modes with various potential mitigations.
The temperature differential can be particularly problematic for rotors in the compressor section as thermal stress can be the limiting parameter in pushing compression ratios higher. Higher compression ratios can improve cycle thermal efficiency and better fuel burn with higher thrust to weight ratios.